This invention is directed to methods and apparatus for precisely producing and precisely assembling a multicomponent article, as for instance an aircraft.
In a manufacturing system it is possible to produce a one piece article with a particular precision by machining the article to that precision. Other articles can be manufactured by first making a die or mold and then forming the article with the die or mold. In using a mold or a die, while the die or mold may be of the required precision, in making the article from the die or mold, precision can and normally is lost.
In manufacturing a multi-component structure, each individual component must be precisely manufactured and then all of the individual components assembled together. Unfortunately, the overall precision of the assembled structure reflects the least precise of its many component parts as well as the precision in joining each of the component parts to each other.
Depending on the article which is being manufactured, the precision required varies. In manufacturing aircraft, the precision of and the alignment of component parts of the aircraft with respect to one another is of a much greater importance than in other manufactured articles. Because of the aero-dynamics of the aircraft, the flight stresses, material fatigue and other similar factors, a very large structure, the aircraft, must be constructed to very exacting tolerances.
An aircraft must be constructed so it is light in weight, but at the same time has sufficient strength to withstand bending, sheer, torsion, and other forces placed on it. To achieve these requirements, modern aircraft are normally constructed utilizing a stressed skin construction.
A stressed skin construction for an aircraft relies upon the strength of the outer skin and attached components and not on strength imparted by internal structural members which might traverse or crisscross through the interior of the aircraft. The stressed skin construction results in a high strength to weight ratio. This is achieved utilizing a thin skin which is reinforced at intervals with bulkheads which are placed transversed to the longitudinal axis of the aircraft and stringers or longerons which are placed along the longitudinal axis of the aircraft. The wings can also be formed utilizing a stressed skin construction, except they can include spars along the longitudinal axis of the wing. In a stressed skin construction it is desirable to achieve an as exact placement and alignment of the individual component parts as possible.
In the past, the design of aircraft structures was done exclusively on two dimensional media, as for instance drafting boards and lofting boards. Drawings of major assemblies and of sub-assemblies were made. A drawing or a loft board was then prepared for each individual structural component specified in the design. While in theory if two drawings are accurately drawn and appropriately dimensioned with respect to one another, the parts made to the specification of these drawings should match. In the practical world, however, this is not necessarily true.
In preparing two separate drawings, inconsistencies do occur, and human judgments made on a 2-dimensional drawing or lofting board do not always prove to be correct in a 3-dimensional world. Further, when one part is joined to another, appropriate alignment of the connecting elements of these parts must be assured. Thus, to produce a modern aircraft it is necessary to use many assembly tools and jigs in order to achieve part alignment. Additionally, assembly jigs require thinking in a further dimension in that they mimic the complementary image of the product and not the image itself, that is the image must fit within the assembly jig.
Typically, in the past, aircraft were formed of metallic components, principally aluminum and aluminum alloys. For more exotic aircraft, as for instance military aircraft, other material such as magnesium and titanium alloys have been used. In aircraft constructed of metallic parts, once the individual parts were formed, during assembly of the aircraft it was possible to make allowances for slight misalignment or precision tolerance differences between parts by using shims or by bending or flexing of certain of the metallic parts, such as thin metallic "skin" sheets. The use of shims or material movement allowed the aircraft assembler to overcome the lack of precision of manufacture of the individual components of the aircraft.
In presently used manufacturing procedures, in order to insure that the component parts of an aircraft do, in fact, fit in alignment with one another, it is necessary to make a family of related master tools including full scale mock-ups or "master models" of major components of the aircraft. Aside from the master model, other master tools include a control (female) master, jig masters, gauging masters, body masters and various templates.
As can be appreciated, many of the master tools for large aircraft, as for instance a Boeing 747, are immense. By their sheer size they are, of necessity, very heavy. Because it is necessary to insure that production aircraft number 999 is the same as production aircraft number 1, these master tools must be stored for the lifetime of the production of the aircraft.
Aside from their size, master tools have further inherent problems. They are easily damaged and can be affected by temperature and humidity. A transfer taken from a master tool at one time may differ from a similar transfer taken from that master tool at a different time because of existing conditions at the time the transfers are taken. Further, in master models, because of their sheer size, a master mold line surface is often independent of the actual engineering design criteria and is extremely difficult to reproduce to the tolerances sought.
With the exception that gauging masters are sometimes independently constructed, the master model must serve as a single mastering source for use in transferring engineering configurations and all other related information, i.e., hole patterns, tool tabs, trim lines, and other criteria to further tools. The master model is used to guarantee interchangeability and replacement, critical dimensional tolerances and match conditions. It is used during fabrication, rework, and verification of other subordinate tools.
In this type of manufacturing system, without the master model, the ability is lost to guarantee dimensional integrity throughout other tools, and in the production parts themselves. The master model is made from engineering data. However, the reproduction of that engineering data in the master model is highly dependent upon the design itself and is very difficult to do.
In the presently utilized manufacturing systems, since further tools are necessary for defining trim lines, hole patterns, and other details which are critical to the final assembly of a highly precision multi-component system such as an aircraft, in order that these further tools give accurate representations of the master model, they have to be "pulled", i.e. molded, directly from the master model or "pulled" second hand from a prior tool which was itself "pulled" from a master model. The master model, thus, served as an image for making further tools, and some of these further tools served as an image for even further tools.
Typically, the control master (a master female mold) is taken off of a component portion of the master model and then utilized to make further image molds of the master model. Inherently, tolerance build-up always occurs when transfers are made from the master model, or from the female control master to any further tool. This is further complicated by the effects of stability of any tool and its response to environmental fluctuation, such as temperature and humidity.
Since in the presently utilized methods of manufacturing aircraft, the master model must first be formed and then the control master pulled from it, the jig master pulled from the control master, the gauging master pulled and etc., etc. prior to the formation of production tooling, the present manufacturing methods for aircraft construction can be considered to be a serial method of manufacturing. By this it is meant that each component tool is dependent on one or more preceding tools and the particular component tool cannot be made until its preceding ancestral tools are first made. By the time production tools are made, many other tools have had to have been created and a lot of time has elapsed. Further in this serial progression of tools, any defects introduced in the lineage of the tools is passed to the descendants further down in the lineage.
In forming an aircraft from a collection of associated parts, assembly jigs are used. Among other things, the assembly jigs are utilized in positioning individual aircraft component parts with respect to one another during final assembly of an aircraft structure. It is necessary, however, to establish the positioning of the assembly jigs themselves to insure that the structures which are positioned with these assembly jigs are accurately positioned. The master tools have also been required for this.
In presently utilized manufacturing systems, an assembly jig rough structure is constructed. A gauging master is rigged in association with the rough structure. Removable locators are designed and manufactured. These locators are then attached to the rough structure in a manner whereby they can later be removed. The locators are designed to "mate" in a known manner with the gauging master and subsequently with the aircraft part or parts which will be assembled on the assembly jig.
Once the locator mounts on the rough structure are established and the locators properly aligned with the gauging master, the locators are removed and the gauging master is also removed. The locators are now reassembled on the rough structure for use in assembly of an aircraft component. When the first aircraft component assembly is completed, the locators are removed to allow for removal of the aircraft component. The locators are then reattached for the next component assembly.
Since the locators are continually being attached and removed from the rough structure, it is necessary to do periodic "cycle checks" or stability audits of the assembly rig. To do this, the gauging master must be re-rigged to the rough structure, the locators mounted, and the total rig checked to see if it is within acceptable tolerance.
Aside from normal cycle checks, the locators can be damaged and are subject to "wear" since they are continually being assembled and disassembled on the rough structure. Anytime a cycle check is performed or a locator is replaced, the particular assembly operation is halted. This not only affects the particular assembly but it can affect related assemblies. A problem with just one locator at a single assembly station can reverberate through many assembly operations.
While cycle checks can be "scheduled" to lessen their impact on other assembly operations, a damaged locator can cause a major disruption of the assembly process because of availability of the gauging master, the time necessary to rig the gauging master on the rough structure and the repositioning of a new locator.
In order to circumvent the necessity of re-rigging the gauging master for a cycle check, a "photogrammetry" process has been tested. In photogrammetry, a glass plate photograph of an assembly is analyzed to verify the position of the assembly. While this process has its merits, it is slow and subject to certain inherent problems such as plate archiving. Further it utilizes a two dimensional process to measure three dimensional spatial parameters. If the photogrammetric process indicates an improper positioning of a locator, the gauging master must be re-rigged to re-align the locator.
As the requirements of aircraft have changed, in order to increase the strength of the aircraft while concurrently decreasing the weight of the aircraft or change other functional properties of the aircraft, new materials have been developed. Presently, composite materials are finding greater use as the construction material for aircraft components. These composite materials offer high strength equal to, or exceeding, that of metallic materials, while at the same time are lighter in weight and have other improved functional properties.
These other functional properties, as for instance increased stiffness and strength of composite parts compared to metallic parts, do not generally lend themselves to the "bend" and "shim" construction techniques utilized in prior metallic aircraft. Since aircraft components formed of composite materials have very rigid structural properties, alignment of adjacent components with one another has become even more critical than in the past. This is a problem not only for the aircraft industry but for other industries, such as automobile manufacturing which is increasingly using composite materials to make body panels and other functional components.
In using composite constructions, it is possible to form a skin panel and stiffeners for the skin panel as a single unit. In prior metal constructions, these normally were formed independently and then joined (using bending and shimming if necessary) with appropriate fasteners such as rivets or bolts. Since the skin panel of a composite component can include other interfacing component members such as spars or longerons formed as an integral part thereon, the criticality of manufacturing parameters including the dimensional precision of these composite structures is even greater than that previously necessary with metal construction techniques. Exact placement and alignment of components with respect to one another in assembling these integrally stiffened composite structures is necessary for a part to be useful. If, in assembling a unified component structure, appropriate parts of the component structure are not exactly placed, the total part is useless and must be scrapped.
Composite components are constructed by laminating structural fibers in appropriate matrixes compatible with these fibers. Fiber glass is a widely used composite system which incorporates glass fibers within an epoxy resin matrix. For formation of aircraft components, more exotic composite systems having improved properties are desirable.
Currently available for use are exotic inorganic materials such as carbon fibers, boron fibers, improved glass fibers, aluminum oxide fibers, inorganic whiskers of different materials and certain organic fibers such as aramides and extended chain polyethylenes. These fibers or whiskers are incorporated as threads, fabrics, mats or the like in appropriate resins, as for instance thermosetting epoxies, polyesters, polyethers, polyimides and bismaleimides or thermoplastic polyamideimines, polyether sulfones, polyetherether ketones, polyphenylene sulfides and other similar polymeric materials.
One construction technique which is common to both metallic aircraft assembly methods and composite components formation is "forming". Parts formed of composite materials are generally formed utilizing molding techniques--using either external molds which are of a complementary shape to a part or an internal mandrel type mold on which the appropriate composite part is built. In either instance the precision of any particular composite part can be no greater than the precision of the mold or tool on which it is made. Highly precise molding techniques are, therefore, necessary in the formation of these composite components. Heretofore, to achieve the necessary precision molds necessary to form composite parts, master tools were also used.
Aside from all of the above, in forming component parts utilizing composite materials, the molds which are utilized must not only serve as a definition surface during lay up of the composite, they also have to -provide support for the composite materials during the curing cycle of these materials. During this curing cycle it is required that the molds retain their shape and stability. The curing is normally affected by applying heat under a vacuum in an autoclave. Thus, not only must the mold be stable in an ambient environment, it must also be stable in the elevated temperature and/or pressure environment of the autoclave in order to maintain the critical shape and dimensions of the composite part during the curing cycle before the composite part has matured into a rigid self-defining structure.
A mold utilized for the formation and curing of a composite part is called a bonding tool. Transfer of trim lines, drilling patterns, and other surface locators to these bonding tools is very labor intensive and exacting. Further, checking aids and feeler gauges must be utilized to insure that these transfers are accurate and are maintained over multiple thermal cycling of the bonding tool in forming repeating units of the composite part thereon.
If an out-of-contour condition is found between a check aid and a bonding tool, not only must the bonding tool be suspect but the stability of the check aid and its master tool must also be suspect. If the check aid does not correlate to the bonding tool and it is then checked against its master tool and it does not correlate with the master tool, the question arises as to whether it is the master tool or the check aid which has an out of contour condition. Further, this is a subjective technique and varies from worker to worker.
Since present tools, including the ones utilized for forming composite parts, are fabricated by a multi-step procedure derived from a master model, the degree of dimensional accuracy achievable is restricted by the material properties of the tools as well as processing techniques. Dimensional accuracy cannot be maintained throughout the multi-step transferring process and, as a result, inconsistency in assembly, production, or intermediate tools has been a significant problem.
The necessity for the "bending" and "shimming" during final assembly of an aircraft can easily be seen because of gradual and/or cumulative tolerance build-up over the many steps necessary to manufacture the individual aircraft component parts. Since bending and shimming are not available in composite manufacturing, these problems are even more dramatic when composite parts are manufactured.
Typically, the presently used master tool or master model is formed from steel, aluminum and plaster. The master facility tool or female control tool taken from the master model normally is a fiber glass tool which is transferred directly from the master model. From this, transfer tools can then be taken which mimic the master model, and from the transfer tool the bonding tool is formed.
The bonding tool is normally a composite tool. These bonding tools in the past have been formed of graphite epoxy structures, fiber glass epoxy structures, ceramic structures, or metal deposited structures. These materials are incorporated as a surface onto supported templates or heavy carrier frames. Problems can result because of thermal mismatch between component parts, residual curing stresses, resin shrinkage, and insufficient process control.
In view of the problems addressed above, it is evident why large precision structures such as aircraft are very costly and require long lead times for their development. The use of composite materials in an aircraft or other structure leads to improved characteristics of that aircraft or structure. However, it also contributes to further manufacturing obstacles in addition to those already inherent in the manufacture of aircraft or other structures.